1. Field of the Invention
The present invention relates generally to turbine blades, and more specifically to cooling of the blade tip.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a hot gas flow is passed through a turbine to drive a compressor and propel an aircraft or, in the case of an industrial gas turbine (IGT) engine, to drive an electric generator and produce electrical energy. The efficiency of the turbine, and therefore the overall engine, can be increased by passing a higher temperature hot gas flow into the turbine.
However, the maximum temperature for the turbine is dependant upon the material properties of the first stage turbine airfoils (stator vanes and rotor blades) and the amount of internal cooling of these airfoils. Turbine airfoil designers attempt to maximize the cooling ability of a given amount of cooling air while at the same time minimizing the amount of cooling air used in order to further increase the engine efficiency. The compressed cooling air used to pass through these hot airfoils is generally bled off from the compressor and thus reduces the engine efficiency.
The rotor blades rotate within the turbine shroud and form a blade tip clearance gap between the blade tip and the inner surface of the shroud in which the hot gas flow passing through the turbine can also leak around. The blade tip leakage reduces the efficiency of the turbine as well as passes hot gas over the tips of the blades and produce hot spots that can lead to blade tip oxidation. The oxidized blade tip can destroy the critical surface shape of the airfoil and lead to decreased performance. Also, oxidized blades have a shortened life and must be replaced. Replacing turbine airfoils requires shutting down the engine and removing the damaged parts. An industrial gas turbine engine requires long running periods of around 48,000 hours. Shutting down an engine prematurely in order to repair damaged parts is costly and results in the loss of use of the engine.
On prior art turbine blade with tip cooling is U.S. Pat. No. 6,916,150 B2 issued to Liang on Jul. 12, 2005 and entitled COOLING SYSTEM FOR A TIP OF A TURBINE BLADE and represented in FIGS. 1 through 5 of this application. FIG. 1 shows the fully assembled turbine blade with the airfoil extending from the platform and root portions of the blade. FIG. 2 shows the airfoil tip portion 11 with a plurality of radial core print-out holes 12 that are typically cast into the airfoil tip and connect to the serpentine coolant passages formed within the airfoil. A tip cap 14 is secured to the blade tip 11 by an adhesive 13 to form the assembled blade. A transient liquid phase (TLP) bonding technique is used to secure the tip cap to the tip of the blade. The tip cap will cover over the core print-out holes 12 in the airfoil. An abrasive layer 15 is applied to the tip cap to promote rubbing with an abradable blade outer air seal (BOAS) surface to form a seal during blade rotation. A great benefit can be obtained with the use of this cooling construction concept for the blade tip cooling design with an abrasive material. as seen in FIGS. 4 and 5, a row of peripheral film cooling holes located on both the pressure side and the suction side of the airfoil just below the tip discharge film cooling air to the tip edge.
A disadvantage of the prior art blade tip cooling design is that the cooling flow distribution and pressure ratio across the film cooling holes for the airfoil pressure and suction sides as well as tip cooling holes are predetermined by the main body serpentine internal cavity pressure. Also, the blade tip region is subject to severe secondary flow field which translates into a large quality of film cooling holes and cooling flow that is required for cooling of the blade tip peripheral.